1. Field of the Invention
This invention is generally directed to satellite attitude determination and control systems and methods, and, more particularly, to satellite attitude determination and control systems and methods that are applicable to spinning spacecraft operations.
2. Description of the Related Art
Transporting a spacecraft from the ground to a destination orbit is an integral and crucial part of any spacecraft mission. For example, to insert a spacecraft into a geosynchronous orbit, a launch vehicle typically injects the spacecraft into a low-altitude parking orbit. The spacecraft then performs transfer orbit operations to transfer the spacecraft from the parking orbit to a destination orbit. The transfer orbit is usually performed by firing a liquid apogee motor (LAM) with the spacecraft spinning around a LAM axis to stabilize the spacecraft and to even the thermal and power conditions, or by firing a combination of LAM and XIP thrusters. Once the spacecraft has completed its transfer orbit, it then may enter in-orbit testing and on-station operation.
From cradle to grave, the spacecraft may go through the following phases of operations: separation, transfer orbit operation (including coasting, spin speed change, reorientation and LAM burn), deployment (including antennas, reflectors, solar wings, radiators), acquisition (including power acquisition and attitude acquisition), in-orbit test (including antenna mapping), on-station operation (including normal pointing, momentum dumping, station keeping and station change), and a deorbiting operation.
Typically, spacecraft, such as communication satellites, use multiple separate sets of sensors and control algorithms for different phases of spaceflight. For example, different sets of sensors and/or control algorithms may be used for attitude determination and control for bi-propellant spinning transfer orbit operations versus those that are used for on-station operations. The use of different sensors, attitude determination, and attitude control methods for spinning transfer orbits and on-station operations, respectively, increases the spacecraft weight, sensor and processor complexity, as well as the development cost for spacecraft attitude determination and control systems.
Spinning transfer orbit operations for spacecraft typically may be performed by ground-assisted attitude determination using a spinning earth sensor and a spinning sun sensor set. The measured leading edge and trailing edge of the earth detected by the earth sensor and the measured TOA (time of arrival) of the sun detected by the sun sensor collected and relayed periodically to a ground station. Typically, at least one orbit pass is dedicated this data collection. A ground orbital operator may then run a ground attitude determination algorithm using these inputs and ephemeris-computed sun and earth positions to determine the spin axis attitude of the spacecraft. This spin axis attitude (the spin phase being still undetermined) is then uploaded to the spacecraft. Next, on-board software may use this spin axis attitude together with the spin phase measured by the spinning sun sensor to complete the 3-axis attitude determination for subsequent spacecraft reorientation or liquid apogee motor (LAM) burn.
Typical spinning transfer orbit operations impose an attitude constraint for the coasting attitude to ensure that the spinning earth and sun sensors will scan through the earth and the sun. This attitude constraint can adversely reduce the power available from solar arrays, or increase the maneuver angle between the coasting attitude and the LAM burn attitude (e.g., by increasing the reorientation maneuver angle for the LAM burn and thus fuel usage).
On-station spacecraft operations typically use different sensors, such as a staring earth sensor assembly (STESA) and a wide field of view (WFOV) sun sensor assembly (SSA), and/or a star tracker for attitude determination. Thus, the sensors used for transfer orbit operations may lie dormant for the entire time that the spacecraft is on station. The number of sensor types used and the number of sensors used, increase the hardware and development cost, increase weight and launch cost, and complicate the mission operation. In addition, some spacecraft have configurations and equipment that may make it difficult in some situations to provide a clear field of view for some sensors, such as, for example, a WFOV SSA, which spans a diamond of about 120×120 deg.
In addition, a wheel-gyro wobble and nutation controller (WGWANC) is typically used for spinning transfer orbit coasting control. A WGWANC can compensate for wobble, capture nutation, and alter spacecraft dynamics by counter-spin or super-spin. However, a WGWANC is very different from the 3-axis stabilized controller typically used for on-station operation. A WGWANC is also susceptible to interact with the fuel slosh dynamics introduced by spacecraft spinning. Fuel slosh is inherently very difficult to model and adds large uncertainty to the WGWANC stability margin. Thus, multiple control types are typically needed for spinning transfer orbit operations versus on-station operations. The use of multiple control types increases the design/analysis/simulation/software/test and other development costs.
This way of operation requires that the spacecraft be in a momentum-conserved dynamic condition (i.e., spacecraft momentum remains unchanging in the earth centered inertial (ECI) frame) during the data collection for the ground-assisted attitude determination. This prohibits the usage of thruster control in the coasting phase while sensor TOA's are being collected. This essentially prohibits the use of thruster control as a back up for the WGWANC controller (useful for spacecraft with WGWANC stability concern due to the aforementioned uncertain fuel slosh interaction).
The present invention is directed to overcoming one or more of the problems or disadvantages associated with the prior art.